(1) Field of the Invention
The present invention relates to microcircuit cooling for the pressure side of a high aspect ratio turbine engine component, such as a turbine blade.
(2) Prior Art
The overall cooling effectiveness is a measure used to determine the cooling characteristics of a particular design. The ideal non-achievable goal is unity, which implies that the metal temperature is the same as the coolant temperature inside an airfoil. The opposite can also occur when the cooling effectiveness is zero implying that the metal temperature is the same as the gas temperature. In that case, the blade material will certainly melt and burn away. In general, existing cooling technology allows the cooling effectiveness to be between 0.5 and 0.6. More advanced technology such as supercooling should be between 0.6 and 0.7. Microcircuit cooling as the most advanced cooling technology in existence today can be made to produce cooling effectiveness higher than 0.7.
FIG. 1 shows a durability map of cooling effectiveness (x-axis) vs. the film effectiveness (y-axis) for different lines of convective efficiency. Placed in the map is a point 10 related to a new advanced serpentine microcircuit shown in FIGS. 2A-2C. This serpentine microcircuit includes a pressure side serpentine circuit 20 and a suction side serpentine circuit 22 embedded in the airfoil walls 24 and 26.
The Table I below provides the dimensionless parameters used to plot the design point in the durability map.
TABLE IOperational Parameters forserpentine microcircuitbeta2.898Tg2581 [F.]Tc1365 [F.]Tm2050 [F.]Tm_bulk1709 [F.]Phi_loc0.437Phi_bulk0.717Tco1640 [F.]Tci1090 [F.]eta_c_loc0.573eta_f0.296Total Cooling Flow3.503%WAE10.8  Legend for Table IBeta = dimensionless heat load parameter or ratio of convective thermal load to external thermal loadPhi_loc = local cooling effectivenessPhi_bulk = bulk cooling effectivenessEta_c_loc = local cooling efficiencyEta_f = film effectivenessTg = gas temperatureTc = coolant temperatureTm = metal temperatureTm_bulk = bulk metal temperatureTco = exit coolant temperatureTci = inlet coolant temperatureWAE = compressor engine flow, pps
It should be noted that the overall cooling effectiveness from the table is 0.717 for a film effectiveness of 0.296 and a convective efficiency (or ability to pick-up heat) of 0.573 (57%). It should also be noted that the corresponding cooling flow for a turbine blade having this cooling microcircuit is 3.5% engine flow. FIG. 3 illustrates the cooling flow distribution for a turbine blade with the serpentine microcircuits of FIGS. 2a-2c embedded in the airfoils walls.
The design shown in FIGS. 2a-2c leads to significant cooling flow reduction. This in turn has positive effects on cycle thermodynamic efficiency, turbine efficiency, rotor inlet temperature impacts, and specific fuel consumption.
It should be noted from FIG. 3 that the flow passing through the pressure side serpentine microcircuit is 1.165% WAE in comparison with 0.428% WAE in the suction side serpentine microcircuit for this arrangement. This represents a 2.7 fold increase in cooling flow relative to the suction side microcircuit. The reason for this increase stems from the fact that the thermal load to the part is considerably higher for the airfoil pressure side. As a result, the height of the microcircuit channel should be a 1.8 fold increase over that of the suction side.
Besides the increased flow requirement on the pressure side, the driving pressure drop potential in terms of source to sink pressures for the pressure side circuit is not as high as that for the suction side circuit. In considering the coolant pressure on the pressure side circuit, FIG. 4 shows that at the end of the third leg, the back flow margin, as a measure of internal to external pressure ratio, is low. As a consequence of this back flow issue, the metal temperature increase beyond that required metal temperature close to the third leg of the pressure side circuit. A remedy is needed to eliminate this problem on the aft pressure side of the airfoil.